Gas turbine engines are internal combustion engines used to provide thrust to an aircraft or to provide power for land-based applications. In general, a gas turbine engine may consist of a fan section and a core engine located downstream of the fan section. In an upstream to downstream direction, the core engine may generally include: 1) a compressor section, which may include a low pressure compressor (LPC) located upstream from a high pressure compressor (HPC), 2) one or more combustors, and 3) a turbine section, which may include a high-pressure turbine (HPT) located upstream from a low-pressure turbine (LPT).
Efforts toward more advanced gas turbine engine designs may be focused on increasing the engine's overall pressure ratio (OPR) which is the ratio of the air pressure at the outlet of the HPC to the air pressure at the inlet of the fan section, as higher OPRs may correlate with higher engine thermodynamic efficiency. To achieve higher OPRs, advanced gas turbine engine designs may have increasing numbers of turbomachinery stages in the compressor section, as each additional turbomachinery stage provides a pressure increase. While effective, this approach may lead to an increased overall engine length and weight, which may present challenges to engine installation and maintenance and may have adverse impacts on rotor dynamics.
Advanced gas turbine engine architectures may also incorporate high-speed LPCs which have higher pressure ratios than traditional low-speed LPCs. However, in contrast with low-speed LPCs which may have radially inwardly curved flowpaths, high-speed LPCs may have straight (or constant radius) flowpaths. As a result, there may be a larger radial offset between the outlet of the LPC and the inlet of the HPC in gas turbine engines incorporating high-speed LPCs. The larger radial offset between the LPC and the HPC may require a longer transition duct between the LPC and the HPC than in earlier engine designs, as curvature in the transition duct may be limited by risks of airflow separation. In particular, the transition duct between the LPC and the HPC may define a roughly “S”-shaped airflow pathway with a first bend at the entrance to the transition duct, and a second bend further downstream prior to entry into the HPC. Pressure gradients may be generated at both an inner wall and an outer wall of the transition duct between the first bend and the second bend as the airflow is turned through the duct, and the magnitude of the pressure gradients may generally increase with increasing curvature or turning angles at the first bend and the second bend. If the pressure gradients are too large, the airflow through the transition duct may separate and reduce the aerodynamic performance of the engine. To keep the pressure gradients within tolerable limits and avoid airflow separation, the curvature of the transition duct may be limited to a certain extent, leading to a more extended structure with increased an axial length with respect to an engine central axis. The curvature constraints on transition ducts may further exacerbate engine length and weight issues in many gas turbine engine designs, particularly in those having high-speed LPCs.
In order to provide transition ducts with shorter axial lengths, current efforts seek to reduce the likelihood of airflow separation in the duct. For example, U.S. Publication Number US 2008/0138197 discloses a transition duct having endwalls with non-axisymmetric perturbations to minimize flow separation and possibly allow for shorter transition ducts with more abrupt curvature. The wall perturbations may take the form of protruding blisters or recessed hollows. While effective, additional enhancements for reducing airflow separation to allow for shorter transition ducts with more pronounced curvature are still wanting.
Clearly, there is a need for improved strategies for reducing the axial lengths of transition ducts in gas turbine engines.